Efficient unmanned aerial vehicle engine simulation process method

Abstract

The invention discloses an efficient unmanned aerial vehicle engine simulation process method, which comprises the following steps: performing fluid domain extraction and grid division on a three-dimensional model to complete data acquisition so as to capture key features of a complex structure of an engine, and in the process of performing fluid domain extraction on the three-dimensional model, performing naming and local encryption so as to complete engine simulation. According to the method, information loss or deviation caused by structural transformation of the obtained fluid model is effectively reduced, it is ensured that the three-dimensional model is highly attached to an actual engine, the modeling precision is remarkably improved, meanwhile, compared with a traditional modeling mode, the starting speed of the whole simulation process is greatly increased, and more time resources are gained for follow-up research.

Description

Technical Field

[0001] The invention belongs to the technical field of engines, and in particular relates to a high-efficiency UAV engine simulation process method. Background Art

[0002] With the rapid development of modern science and technology, UAV engine performance optimization is one of the core research directions in this technical field. In order to break through the limitations of traditional experimental methods, simulation technology is introduced into the field of UAV engine performance research.

[0003] In the early stage of simulation, how to accurately convert the engine structure into a fluid calculation model to improve modeling efficiency while ensuring model accuracy is one of the difficult problems that need to be overcome. When generating a fluid model from a three-dimensional model, due to the complex internal structure of the engine, there is a lack of clear standards and methods in the process of extracting the fluid domain. When assembling the models of various components (such as compressors, combustion chambers, tail nozzles, etc.) to build a complete engine simulation model, the process is unclear and the dilemma is prominent. The combination conditions such as the connection method of each component, the matching of interface parameters, and the collaborative working mechanism under different working conditions are incomplete, which makes the model assembly prone to errors or unreasonable, affecting the accuracy and reliability of the simulation model. When simulating the whole engine, the boundary layer on the surface of the engine causes a significant change in the internal airflow, and the thickness and flow characteristics of the boundary layer are affected by many factors. In addition, the setting of the boundary conditions of the engine simulation calculation (such as intake and exhaust conditions, etc.) is extremely critical. Inaccurate or unstable boundary conditions can cause numerical instability, convergence difficulties, and even errors in the calculation results during the calculation process, and it is impossible to effectively use simulation calculations to deeply study engine performance and carry out optimization design.

[0004] There are many technical difficulties in the existing simulation process that need to be solved. From the construction of three-dimensional models, the generation of fluid models to the assembly, there is a lack of clear and standardized operating procedures and technical specifications, and the connection and transition between the various links are unclear, which increases the uncertainty of the entire simulation process. Especially when it comes to the combination of key engine components such as compressors, combustion chambers, tail nozzles, etc., the combination conditions are not defined accurately and completely, and it is impossible to clarify the adaptation parameters and collaborative working modes of each component under different working conditions and performance requirements, resulting in obvious defects in the accuracy and reliability of the combined engine model. In addition, during the simulation calculation of the entire engine, due to the complex characteristics of the boundary layer and the difficulty in accurately setting and controlling the boundary conditions, the calculation process is often unstable, and the calculation results have large deviations and fluctuations, which cannot provide a solid and reliable data foundation and decision-making basis for the in-depth analysis and optimization design of engine performance. Summary of the invention

[0005] In order to solve the above problems existing in the prior art, the present invention provides an efficient UAV engine simulation process method. The technical problem to be solved by the present invention is achieved through the following technical solutions:

[0006] The present invention provides an efficient UAV engine simulation process method, the method comprising:

[0007] Establishing a three-dimensional model of the engine; the three-dimensional model includes: a compressor, a combustion chamber and a tail nozzle;

[0008] Extracting the fluid domain of the compressor, combustion chamber and tail nozzle in the three-dimensional model to obtain first fluid models corresponding to the compressor, combustion chamber and tail nozzle respectively; naming each face in the first fluid model, locally encrypting the preset position in the first fluid model to obtain a second fluid model;

[0009] Meshing the second fluid models corresponding to the compressor, the combustion chamber and the tail nozzle respectively to obtain corresponding meshes;

[0010] Perform simulation calculations on the grids corresponding to the compressor, combustion chamber and tail nozzle to obtain the boundary conditions corresponding to the compressor, combustion chamber and tail nozzle;

[0011] Insert boundary conditions corresponding to the compressor, combustion chamber and tail nozzle to obtain corresponding balance data;

[0012] Data is inserted into the grids corresponding to the compressor, the combustion chamber and the tail nozzle according to the calculated convergence field in the balance data to obtain the simulation result of the engine.

[0013] In one embodiment of the present invention, fluid domain extraction is performed on the compressor, combustion chamber and tail nozzle in the three-dimensional model to obtain first fluid models corresponding to the compressor, combustion chamber and tail nozzle, respectively, including:

[0014] Deleting the parts irrelevant to the fluid in the three-dimensional model to obtain a processed three-dimensional model;

[0015] The processed 3D model is subjected to fluid region extraction through volume extraction to obtain the fluid domains corresponding to the compressor, combustion chamber and tail nozzle;

[0016] The geometric defects in the fluid domains corresponding to the compressor, the combustion chamber and the tail nozzle are repaired to obtain the first fluid models corresponding to the compressor, the combustion chamber and the tail nozzle respectively.

[0017] In one embodiment of the present invention, naming each face in the first fluid model includes:

[0018] The compressor inlet, compressor outlet, combustion chamber inlet, combustion chamber outlet, tail nozzle inlet and tail nozzle outlet in the first fluid model are named; wherein,

[0019] In the first fluid model, the parts of the rotor that need to rotate, the rotor casing, the rotor blades and the inner wall are named separately.

[0020] In one embodiment of the present invention, the preset positions include: rotor and stator blades in the compressor and tail nozzle.

[0021] In one embodiment of the present invention, the second fluid models corresponding to the compressor, the combustion chamber and the tail nozzle are respectively meshed to obtain corresponding meshes, including:

[0022] Select hybrid grid as the grid type and set the boundary layer of the grid;

[0023] Dividing the preset key areas in the second fluid model corresponding to the compressor, the combustion chamber and the tail nozzle;

[0024] Discretize the second fluid model corresponding to the compressor, combustion chamber and tail nozzle according to the grid type, grid boundary layer and preset key area, and generate the corresponding primary grid;

[0025] The quality of the first-level grid is checked according to the preset indicators, and the first-level grid that meets the preset conditions is used as the corresponding grid for the compressor, combustion chamber and tail nozzle.

[0026] In one embodiment of the present invention, the preset key areas include: valves, fuel injection nozzles and bends.

[0027] In one embodiment of the present invention, the preset indicators include: orthogonality, aspect ratio and Jacobian matrix.

[0028] In one embodiment of the present invention, the boundary conditions include: inlet boundary, outlet boundary, wall condition, periodic boundary and special boundary.

[0029] In one embodiment of the present invention, during the simulation calculation of the compressor and the tail nozzle, the rotation direction of the compressor rotor is set to the intake direction, the compressor rotor is set to rotate clockwise, and the tail nozzle rotor is set to rotate counterclockwise.

[0030] In one embodiment of the present invention, during the process of inserting boundary conditions corresponding to the compressor, the combustion chamber and the tail nozzle,

[0031] After the compressor calculation converges, the compressor outlet boundary conditions are derived as the combustion chamber inlet settings;

[0032] After the combustion chamber calculation converges, the combustion chamber outlet boundary conditions are derived as the tail nozzle inlet settings;

[0033] The convergence data corresponding to the compressor, combustion chamber and tail nozzle are inserted as conditions for rapid equilibrium.

[0034] Beneficial effects of the present invention:

[0035] In the solution provided by the present invention, data collection is completed by performing fluid domain extraction and grid division on the three-dimensional model, thereby capturing the key features of the complex structure of the engine. In the process of fluid domain extraction on the three-dimensional model, the obtained fluid model is effectively reduced through naming and local encryption, ensuring that the three-dimensional model is highly consistent with the actual engine. While significantly improving the modeling accuracy, compared with the traditional modeling method, the startup speed of the entire simulation process is greatly improved, thereby gaining more time resources for subsequent research. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 A schematic diagram of the steps of an efficient UAV engine simulation process method provided by an embodiment of the present invention;

[0037] Figure 2 A schematic diagram of a three-dimensional model of an engine in an efficient UAV engine simulation process method provided by an embodiment of the present invention;

[0038] Figure 3 A schematic diagram of a fluid model of an engine in an efficient UAV engine simulation process method provided by an embodiment of the present invention;

[0039] Figure 4 A schematic diagram of a grid of an engine in an efficient UAV engine simulation process method provided by an embodiment of the present invention;

[0040] Figure 5 A schematic diagram of a compressor part in a high-efficiency UAV engine simulation process method provided by an embodiment of the present invention;

[0041] Figure 6 A schematic diagram of the stability of mass flow rate in a high-efficiency UAV engine simulation process method provided by an embodiment of the present invention;

[0042] Figure 7 A schematic diagram of the flow conditions of a compressor and a combustion chamber in a high-efficiency UAV engine simulation process method provided by an embodiment of the present invention;

[0043] Figure 8 A schematic diagram of the integration of an engine as a whole in an efficient UAV engine simulation process method provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0044] The present invention is further described in detail below with reference to specific embodiments, but the embodiments of the present invention are not limited thereto.

[0045] The embodiment of the present invention provides an efficient UAV engine simulation process method, such as Figure 1 As shown, this may include:

[0046] S1, build a three-dimensional model of the engine; the three-dimensional model includes: compressor, combustion chamber and tail nozzle.

[0047] Specifically, the three-dimensional model of the engine is as follows: Figure 2 As shown, it is divided into three areas, from top to bottom: compressor, combustion chamber and tail nozzle.

[0048] Building a 3D model of the engine can include:

[0049] The solid model of the engine is obtained as a three-dimensional model using CAD software or CAE pre-processing tools.

[0050] CAD software may include NX. CAE pre-processing tools may include: ANSYS SpaceClaim or DesignModeler.

[0051] In the process of building the three-dimensional model of the engine, it is necessary to ensure that the model is a closed entity without gaps or overlapping surfaces.

[0052] S2, extracting the fluid domain of the compressor, combustion chamber and tail nozzle in the three-dimensional model to obtain first fluid models corresponding to the compressor, combustion chamber and tail nozzle respectively; naming each face in the first fluid model, locally encrypting the preset position in the first fluid model to obtain a second fluid model, which may include:

[0053] S21, deleting the components irrelevant to the fluid in the three-dimensional model to obtain a processed three-dimensional model; wherein the components irrelevant to the fluid may include bolts and brackets.

[0054] S22, extracting the fluid region of the processed three-dimensional model through volume extraction to obtain the fluid domains corresponding to the compressor, the combustion chamber and the tail nozzle.

[0055] For example, you can subtract the piston solid from the engine cylinder cavity to obtain the fluid domain of the combustion chamber.

[0056] S23, repairing geometric defects in the fluid domains corresponding to the compressor, the combustion chamber, and the tail nozzle, and obtaining first fluid models corresponding to the compressor, the combustion chamber, and the tail nozzle respectively.

[0057] Specifically, the geometric cleaning tool can be used to fill small holes, repair overlapping surfaces, and simplify small features (such as chamfers) in the geometric defects in the fluid domain to obtain the following Figure 3The first fluid models corresponding to the compressor, combustion chamber and tail nozzle are shown to ensure the continuity of the fluid domain surface and avoid subsequent meshing failure.

[0058] S24, naming each face in the first fluid model, and locally encrypting a preset position in the first fluid model to obtain a second fluid model.

[0059] In S24, naming each surface in the first fluid model may include:

[0060] The compressor inlet, compressor outlet, combustion chamber inlet, combustion chamber outlet, tail nozzle inlet and tail nozzle outlet in the first fluid model are named; wherein,

[0061] In the first fluid model, the parts of the rotor that need to rotate, the rotor casing, the rotor blades and the inner wall are named separately.

[0062] The preset positions in the first fluid model may include: rotors and stator blades in the compressor and the tail nozzle.

[0063] It can be understood that by locally encrypting the preset position in the first fluid model, the obtained second fluid model can meet the accuracy of subsequent grid division and meet the requirements of simulation calculation.

[0064] S3, mesh the second fluid models corresponding to the compressor, combustion chamber and tail nozzle respectively to obtain corresponding meshes, such as Figure 4 As shown, it may include:

[0065] S31, select hybrid grid as the grid type of the grid, and set the boundary layer of the grid; wherein, the hybrid grid can be set to use prism layers in key areas (such as boundary layers) and tetrahedrons in other areas; setting the boundary layer of the grid can include: generating boundary layer grids near the wall surface, ensuring that the first layer of grids meets the requirements of the turbulence model, the number of layers can be set to 5-15 layers, and the growth ratio can be set to 1.2-1.5.

[0066] S32, dividing the preset key areas in the second fluid model corresponding to the compressor, the combustion chamber and the tail nozzle respectively; wherein the preset key areas may include: valves, fuel injectors and bends.

[0067] S33, discretizing the second fluid models corresponding to the compressor, combustion chamber and tail nozzle respectively according to the grid type, the boundary layer of the grid and the preset key area, and generating the corresponding first-level grid; wherein, in the process of generating the corresponding first-level grid, the preset key area can be locally encrypted using the size function or volume control to control the grid size.

[0068] S34, checking the quality of the first-level grid according to the preset index, and using the first-level grid that meets the preset conditions as the grid corresponding to the compressor, the combustion chamber and the tail nozzle.

[0069] Specifically, the preset indicators include: orthogonality, aspect ratio and Jacobian matrix.

[0070] When checking the quality of the primary grid according to preset indicators, it can be set that the orthogonality needs to be greater than 30°, the aspect ratio needs to be less than 5, and the Jacobian matrix needs to be greater than 0.3.

[0071] The key to connecting various regions in the engine model lies in the processing of the interface and the alternation of data. First, the mesh at the interface must be consistent. For example, the mesh of the compressor is Figure 5 As shown, the number and quality of meshes in the bottom area are required to be similar to the combustion chamber part, followed by data transmission of the interface, and the compressor inlet conditions are determined by the experimental inlet conditions.

[0072] S4, simulate and calculate the grids corresponding to the compressor, combustion chamber and tail nozzle to obtain the boundary conditions corresponding to the compressor, combustion chamber and tail nozzle; wherein the boundary conditions include: inlet boundary, outlet boundary, wall condition, periodic boundary and special boundary.

[0073] Inlet boundaries can include:

[0074] Speed ​​entry: specify speed magnitude and direction;

[0075] Mass flow inlet: suitable for compressive flow;

[0076] Total Pressure Inlet: For compressible flow.

[0077] Export boundaries can include:

[0078] Pressure outlet: set static pressure or ambient pressure;

[0079] Mass flow outlet: suitable for reflux control.

[0080] Wall conditions can include:

[0081] No-slip or slip conditions, adiabatic walls or heat transfer walls, and roughness settings.

[0082] Periodic boundaries can include: If the geometry is symmetric. Set periodic conditions to reduce the amount of calculations.

[0083] Special boundaries can include rotating parts (such as turbine blades): use a moving reference frame or a sliding network.

[0084] In the process of simulating the compressor and tail nozzle, the rotation direction of the compressor rotor is set to the intake direction, the compressor rotor is set to rotate clockwise, and the tail nozzle rotor is set to rotate counterclockwise. While the rotor is rotating, the outer wall of the rotor should be kept stationary relative to the rotor.

[0085] Specifically, the compressor is simulated and calculated, the mass flow inlet is set at the inlet of the air inlet, the speed of the rotor is set according to the experimental data, and the pressure outlet is set according to the compression ratio at the stator outlet; the combustion chamber imports the compressor outlet pressure data based on the compressor calculation, the fuel inlet of the combustion chamber is set according to the experimental data, and the outlet is set with a pressure outlet; the tail nozzle imports the combustion chamber outlet pressure data based on the combustion chamber calculation. During the calculation process, the diagram of the stability of the mass flow rate can be found in Figure 6 As shown, the inlet and outlet mass flow rate errors are monitored, and when the residual tends to converge and the mass flow rate error is within 15%, it means that the calculation condition has converged.

[0086] S5, insert boundary conditions corresponding to the compressor, combustion chamber and tail nozzle to obtain corresponding balance data.

[0087] Specifically, when the flow field of the compressor part tends to be stable, the outlet pressure is recorded, and the Profile pressure attribute of the outlet part is derived to complete the initial pressure initialization of the combustion chamber inlet air end. The fuel inlet is assigned a fuel mass flow rate fraction according to the situation given by the experimental department.

[0088] When the combustion chamber calculation is stable and the compressor part is the same, first record the combustion chamber outlet pressure and mass flow rate. Then insert the combustion chamber flow field as data into the compressor part to make the two reach convergence as soon as possible. For a schematic diagram of the flow conditions of the compressor and combustion chamber, please refer to Figure 7 As shown in the figure, it can be seen that after the two are connected, their respective stable flow fields and data can circulate, react and flow with each other.

[0089] In the process of inserting boundary conditions corresponding to the compressor, combustion chamber and tail nozzle,

[0090] After the compressor calculation converges, the compressor outlet boundary conditions are derived as the combustion chamber inlet settings;

[0091] After the combustion chamber calculation converges, the combustion chamber outlet boundary conditions are derived as the tail nozzle inlet settings;

[0092] The convergence data corresponding to the compressor, combustion chamber and tail nozzle are inserted as conditions for rapid equilibrium.

[0093] The tail nozzle part uses the combustion chamber end outlet condition, which is the same as the combination of the compressor and the combustion chamber. First, the tail nozzle inlet data is selected as: the combustion chamber outlet pressure and mass flow rate after the compressor and the combustion chamber are calculated together. Set the tail nozzle outlet boundary conditions, and after the tail nozzle part is calculated separately, perform the overall simulation calculation with the compressor and the combustion chamber.

[0094] S6, inserting data into the grids corresponding to the compressor, combustion chamber and tail nozzle according to the calculated convergence field in the balance data to obtain the simulation results of the engine.

[0095] For a schematic diagram of the complete engine assembly, see Figure 8 As shown, Figure 8 The variation of oxygen mass fraction in the whole engine under the condition of continuous operation with steady flow field insertion and boundary conditions is given. The combination of the three parts clearly shows the transferability of oxygen in the combustion chamber and tail nozzle.

[0096] The embodiment of the present invention completes data collection by performing fluid domain extraction and grid division on the three-dimensional model, thereby capturing the key features of the complex structure of the engine. In the process of fluid domain extraction on the three-dimensional model, the obtained fluid model is effectively reduced through naming and local encryption, ensuring that the three-dimensional model is highly consistent with the actual engine. While significantly improving the modeling accuracy, compared with the traditional modeling method, the startup speed of the entire simulation process is greatly improved, thereby gaining more time resources for subsequent research.

[0097] It should be noted that in the description of the present invention, it should be understood that the terms "first" and "second" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, the meaning of "plurality" is two or more, unless otherwise clearly and specifically defined.

[0098] The above description is only a preferred embodiment of the present invention and is not intended to limit the protection scope of the present invention. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention are included in the protection scope of the present invention.

Claims

1. An efficient UAV engine simulation process method, characterized in that: include: Build a 3D model of the engine; The three-dimensional model includes: a compressor, a combustion chamber and a tail nozzle; Extracting the fluid domain of the compressor, combustion chamber and tail nozzle in the three-dimensional model to obtain first fluid models corresponding to the compressor, combustion chamber and tail nozzle respectively; naming each face in the first fluid model, locally encrypting the preset position in the first fluid model to obtain a second fluid model; Meshing the second fluid models corresponding to the compressor, the combustion chamber and the tail nozzle respectively to obtain corresponding meshes; Perform simulation calculations on the grids corresponding to the compressor, combustion chamber and tail nozzle to obtain the boundary conditions corresponding to the compressor, combustion chamber and tail nozzle; Insert boundary conditions corresponding to the compressor, combustion chamber and tail nozzle to obtain corresponding balance data; Data is inserted into the grids corresponding to the compressor, the combustion chamber and the tail nozzle according to the calculated convergence field in the balance data to obtain the simulation result of the engine.

2. The method for simulating an efficient UAV engine according to claim 1, characterized in that: Perform fluid domain extraction on the compressor, combustion chamber and tail nozzle in the three-dimensional model to obtain first fluid models corresponding to the compressor, combustion chamber and tail nozzle, respectively, including: Deleting the parts irrelevant to the fluid in the three-dimensional model to obtain a processed three-dimensional model; The processed 3D model is subjected to fluid region extraction through volume extraction to obtain the fluid domains corresponding to the compressor, combustion chamber and tail nozzle; The geometric defects in the fluid domains corresponding to the compressor, the combustion chamber and the tail nozzle are repaired to obtain the first fluid models corresponding to the compressor, the combustion chamber and the tail nozzle respectively.

3. The method for simulating an efficient UAV engine according to claim 1 is characterized in that: The step of naming each surface in the first fluid model includes: The compressor inlet, compressor outlet, combustion chamber inlet, combustion chamber outlet, tail nozzle inlet and tail nozzle outlet in the first fluid model are named; wherein, In the first fluid model, the parts of the rotor that need to rotate, the rotor casing, the rotor blades and the inner wall are named separately.

4. The method for simulating an efficient UAV engine according to claim 1, characterized in that: The preset positions include: rotor and stator blades in the compressor and tail nozzle.

5. The method for simulating an efficient UAV engine according to claim 1, characterized in that: The second fluid models corresponding to the compressor, the combustion chamber and the tail nozzle are respectively meshed to obtain corresponding meshes, including: Select hybrid grid as the grid type and set the boundary layer of the grid; Dividing the preset key areas in the second fluid model corresponding to the compressor, the combustion chamber and the tail nozzle; Discretize the second fluid model corresponding to the compressor, combustion chamber and tail nozzle according to the grid type, grid boundary layer and preset key area, and generate the corresponding primary grid; The quality of the first-level grid is checked according to the preset indicators, and the first-level grid that meets the preset conditions is used as the corresponding grid for the compressor, combustion chamber and tail nozzle.

6. The high-efficiency UAV engine simulation process method according to claim 5 is characterized in that: The preset key areas include: valves, fuel injectors and bends.

7. The method for simulating an efficient UAV engine according to claim 5, characterized in that: The preset indicators include: orthogonality, aspect ratio and Jacobian matrix.

8. The method for simulating an efficient UAV engine according to claim 1, characterized in that: The boundary conditions include: inlet boundary, outlet boundary, wall condition, periodic boundary and special boundary.

9. The high-efficiency UAV engine simulation process method according to claim 1 is characterized in that: In the process of simulating the compressor and the tail nozzle, the rotation direction of the compressor rotor is set to the intake direction, the compressor rotor is set to rotate clockwise, and the tail nozzle rotor is set to rotate counterclockwise.

10. The method for simulating an efficient UAV engine according to claim 1, characterized in that: In the process of inserting boundary conditions corresponding to the compressor, combustion chamber and tail nozzle, After the compressor calculation converges, the compressor outlet boundary conditions are derived as the combustion chamber inlet settings; After the combustion chamber calculation converges, the combustion chamber outlet boundary conditions are derived as the tail nozzle inlet settings; The convergence data corresponding to the compressor, combustion chamber and tail nozzle are inserted as conditions for rapid equilibrium.

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